Electrically caged or restrained, two-axis gyros are known. Such gyros have a gyro rotor, which rotates about a spin axis. The spin axis is angularly movable, relative to a housing, with two degrees of freedom about two mutually orthogonal input axes, each of which is orthogonal to the spin axis. A pick-off is arranged on each of the input axes and responds to deflections of the gyro rotor about this input axis. And a torquer is arranged on each of the input axes to exert a torque on the gyro rotor about the input axis. The signal from each of the two pick-offs, after amplification, is applied to the torquer arranged on the respective other input axis. In accordance with the gyro laws, the torque exerted by the torquer counteracts the deflection causing the signal from the pick-off. Thereby the gyro rotor is caged or restrained to a central position about the two input axes. Such electrically caged, two-axis gyros are described in DE-B-29 03 282, in DE-A-30 33 281 and EP-A-0 251 157.
Such gyros are subjected to systematic errors. From DE-C2-29 22 411 it is known to at least partly eliminate such systematic errors by making at least two measurements with the gyro at two positions of the gyro which are angularly offset by 180.degree. about one of the input axes. It is also known to measure with a gyro at three angular positions which are angularly offset by 90.degree. from each other (DE-A-29 22 412).
From DE-C2-31 43 527 and US-A-4 461 089 it is further known, in order to eliminate certain gyro errors, to measure in two positions of the gyro angularly spaced by 180.degree. about the spin axis of the gyro, and to form the difference of the measured values. In this case, the spin axis is vertical. A north angle is determined from the ratio of the differences of the signals applied to the two torquers.
EP-A2-0 263 777 shows an integrated, redundant reference system for flight control and for the generation of heading and attitude information. A plurality of electrically caged, two-axis gyros are arranged aircraft-fixed in such a way that they provide angular rate information redundantly. A plurality of accelerometers provide correspondingly redundant acceleration information. Signal processing means are provided by which erroneous angular rate and acceleation information is eliminated. The thus failure-compensated angular rate and acceleration information provides stabilizing signals for the flight controller. Furthermore, heading and attitude information is generated from this failure-compensated angular rate and acceleration information.
With the reference system as disclosed in EP-A2-0 263 777, a coarse alignment and subsequently a fine alignment of the system take place. Velocity increment signals from the accelerometers are transformed from the system-fixed coordinate system into an earth-fixed coordinate system in accordance with a directional cosine matrix. The transformed velocity increment signals are applied to a Kalman filter which models the influence of the rotation of the earth on the angular rate sensors and generates actuating signals independent of the rotation of the earth for the heading and attitude correction, when the coordinate system defined by the directional cosine matrix deviates from the earth-fixed coordinate system.
DE-A-31 01 828 relates to a method for determining the zero error of a single-axis rate gyro. In order to determine the zero error, and in particular to determine the angle between a reference axis and geographic north, the output signals of the rate gyro are measured at different rotary speeds of the gyro rotor. The sum of all zero errors are determined by computational combination of the measuring results.
DE-C2-32 13 720 describes a dynamically tuned gimbal suspension with two degrees of freedom for a gyro rotor.